1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a ceramic core used to make an industrial gas turbine first or second stage blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
Conventional gas turbine engines include a compressor, a combustor, and a turbine. Air flows axially through the sections of the engine. As is well known in the art, air compressed in the compressor is mixed with fuel which is burned in the combustor and expanded in the turbine, thereby rotating the turbine and driving the compressor. The turbine components are subjected to a hostile environment characterized by the extremely high temperatures and pressures of the hot products of combustion that enter the turbine. In order to withstand repetitive thermal cycling in, such a hot environment structural integrity and cooling of the turbine airfoils must be optimized.
Cooling schemes for airfoils have become very sophisticated in modern engines. The airfoils include intricate internal cooling passages that extend radially within the very thin airfoil. The radial passages are frequently connected by a plurality of small crossover holes to allow the flow of cooling air between the passages. Fabrication of airfoils with such small internal features necessitates a complicated multi-step manufacturing process.
A problem with the current manufacturing process is that it is characterized by relatively low yields. The main reason for the low yields is that during the manufacturing process of airfoils, a ceramic core that defines the cooling passages of the airfoil often either breaks or fractures. There are a number of factors that contribute to such a high percentage of ceramic cores becoming damaged. First, ceramic, in general, is a brittle material. Second, the airfoils are very thin and subsequently, the cores are very thin. Finally, the small crossover holes in the airfoil result in narrow fingers in the core that is easily broken under load.
Another major factor contributing to cores being damaged during airfoil fabrication is that the fragile cores are handled repeatedly, undergoing many manufacturing processes, thereby increasing the chances for the core to break. In such processes, the core is first manually removed from a die and can be easily broken during handling. Subsequently, the core is secured within a mold and pressurized wax is injected into the mold around the core. As the pressurized wax is injected around the core, the core is subjected to shearing, bending and torsion loads that may either crack or break the core. The wax mold, with the core secured inside, is then dipped into a slurry to form layers of coating or a “shell”. The wax is then melted out from the shell, forming a mold with the core secured therein. The shell, with the core secured therein, is subsequently heated. The heating process of the shell with the core results in different rates of expansion of the ceramic core and the shell. The difference in growth of the shell and the core frequently results in the core being fractured or broken since the shell generally expands at a faster rate than the core, thereby stretching the core and breaking it. The next step in the manufacturing process is injecting molten metal into the shell with the core secured therein. As the molten metal is poured into the shell, it may have non-uniform flow, causing shear, bending, and torsion loads on the core. As the molten metal solidifies, the core is then chemically removed from the airfoil. Once the core is removed, the area occupied by the core becomes the internal cavity for cooling air to pass through within the airfoil.
Fractures or breakage of the core during the manufacturing process is frequently detected only after the part is completed. Even a hairline fracture in the core developed at any stage of the manufacturing process will undermine the integrity of an airfoil and result in necessary of scrapping the finished part. One financial disadvantage of obtaining a low yield of good parts is that the effective cost of each usable airfoil is very high.
Another drawback is that the fragile nature of the ceramic cores results in production constraints that limit more optimal cooling schemes. In many instances it may be more advantageous for the airfoil cooling and engine efficiency to have smaller crossover holes or more intricate geometric features. However, more intricate cooling passages are not practical at this time, since the current manufacturing process already yields an insufficiently small number of usable airfoils and has a high percentage of ceramic cores being damaged. More intricate cooling schemes would result in even lower manufacturing yields and even higher, cost per airfoil. Thus, there is a great need to improve manufacturability of the gas turbine engine airfoils to reduce the cost of each airfoil as well as to improve cooling schemes therefor.
The prior art U.S. Pat. No. 5,599,166 issued to Deptowicz et al on Feb. 4, 1997 and entitled CORE FOR FABRICATION OF GAS TURBINE ENGINE AIRFOILS discloses a ceramic core for making a turbine blade and the turbine blade made from the core. The turbine blade in this patent is a small blade of about 2 inches in spanwise length that is used in an aero gas turbine engine. The Deptowicz et al invention makes use of fingers 148, 166 and 172 in the core that corresponds to the crossover holes (also called metering holes) and provides the structural rigidity to the core during the casting process. One problem with using the teaching of the Deptowicz et al disclosure is that scaling of the core for use to make a large turbine blade in an industrial gas turbine engine will result in metering holes too large, resulting in a large cooling air flow through the trailing edge metering holes.
The Green et al U.S. Pat. No. 5,403,159 issued on Apr. 4, 1995 and entitled COOLABLE AIRFOIL STRUCTURE discloses a turbine airfoil used in an aircraft, the airfoil having a trailing edge cooling circuit with a double impingement cooling hole arrangement. The impingement holes are formed in a pair of spanwise extending ribs 106 and 108 with cooling holes 112 formed therein. A cooling air passage exists between the ends of the ribs and the interior surface of the airfoil. The length between the first rib 106 to the surface 22 is ½ the length of the second rib 108 to the surface 22.
Two prior art patents issued to Liang (the inventor of this present invention) U.S. Pat. No. 5,975,851 issued on Nov. 2, 1999 and entitled TUURBINE BLADE WITH TRAILING EDGE ROOT SECTION COOLING and U.S. Pat. No. 6,139,269 issued on Oct. 31, 2000 and entitled TURBINE. BLADE WITH MULTI-PASS COOLING AND COOLING AIR ADDITION both disclose a first stage turbine blade used in an aircraft gas turbine engine. Both blades include a trailing edge cooling circuit with double impingement holes formed in the ribs that extend between a tip pedestal at the top and a platform at the bottom. Impingement holes are formed in the two ribs, and root impingement holes (60 in this patent) are formed between the rib and the platform. The cooling flow area of the holes 60 and 61 are equal to ½ the cooling flow area of the impingement holes 58 and 59.
The known methods of forming a turbine blade used in an aero engine will not work for the same design scaled to the size of a first or second stage turbine blade used in an industrial gas turbine (IGT) engine. The first stage turbine blade in an aero engine is from one to two inches in length, while the same stage blade in an IGT engine is five to ten inches in length. For one thing, the cooling flow path is much longer in the IGT blade than in the aero blade. Also, the ceramic core uses very small impingement cooling holes in the aero blade. If the impingement holes in the ceramic core used to make the aero blade were scaled up to make an IGT blade, the holes would be too large and discharge more cooling air than needed. Thus, a twin impingement cooling hole arrangement shown in the Green '159 patent; the Liang '851 patent; and the Liang '269 patent would not be practical in a scaled up version for an IGT blade.
The object of the present invention is to provide for a ceramic core used to form a turbine blade used in an industrial gas turbine engine that will be stronger than those of the cited prior art such that less defective blades are cast.
Another object of the present invention is to provide for a turbine blade for use in an industrial gas turbine engine with a more efficient cooling circuit for the blade.
Still another object of the present invention is to provide for a turbine blade for use in an industrial gas turbine engine with improved means of controlling and directing internal cooling air within the blade.